This invention relates generally to gas turbine engines and more particularly to film cooled combustor liners used in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Combustors used in aircraft engines typically include inner and outer combustor liners. The liners contain the combustion process and facilitate the distribution of air to the various combustor zones in prescribed amounts.
Because they are exposed to intense heat generated by the combustion process, combustor liners are cooled to meet life expectancy requirements. Liner cooling is commonly provided by diverting a portion of the compressed air (which is relatively cool) and causing it to flow over the outer surfaces of the liners. In addition, a thin layer of cooling air is provided along the combustion side of the liners by directing cooling air flow through cooling holes formed in the liners. This technique, referred to as film cooling, reduces the overall thermal load on the liners because the mass flow through the cooling holes dilutes the hot combustion gas next to the liner surfaces, and the flow through the holes provides convective cooling of the liner walls. There are two basic types of liners that employ film cooling: multi-hole cooled liners and slot cooled liners.
Multi-hole cooled liners include a large number of angled, very small cooling holes formed through the liners. Compressor air passes through the cooling holes to produce the film of cooling air on the combustion side of the liners. The cooling holes are generally distributed over the whole liner so as to provide a constant replenishing of the cooling film along the entire length of the liner. Slot cooled liners include a plurality of connected panel sections with a bump or nugget formed on the forward end of each panel section. An axially oriented slot is formed on the hot gas side surface of each panel section at the nugget, and a circumferentially disposed row of cooling holes is formed in the nugget. Compressor air passes through the cooling holes to produce the film of cooling air on the hot gas side surface of the panel section. Thus, the cooling film is replenished at each slot.
The distribution of air is accomplished through so-called dilution holes in the liners. The dilution holes introduce jets of air to the primary and secondary zones of the combustion chamber. The dilution air quenches the flames so as to control the gas temperature to which the turbine hardware downstream of the combustor will be exposed. The quenching also reduces the level of NOx emissions in the engine exhaust. It is common that different dilution holes have different sizes, depending on the amount of dilution air needed in a particular area of the combustion chamber. For slot cooled liners, dilution holes are currently located in a panel with the hole centers being axially aligned in a circumferential row, which is typically positioned mid-span of the panel. When there are different size dilution holes with this arrangement, the aft edges of the smaller diameter holes are located farther from the downstream slot that will replenish the cooling film.
The wake produced by the influx of air through the dilution holes will disrupt the cooling film. Hot combustion gases can become entrained in these wakes and significantly increase liner metal temperatures. Thus, while film cooling of combustor liners is generally quite effective, the presence of dilution holes can result in hot spots being formed immediately downstream thereof. As a result, current combustor liners can experience reduced low cycle fatigue life, increased oxidation rates of the substrate resulting in spallation of the thermal barrier coating, and accelerated creep of the slot overhangs.
Accordingly, there is a need for a slot cooled combustor liner in which the effect of the dilution holes on the film cooling effectiveness is minimized.
The above-mentioned need is met by the present invention which provides a gas turbine combustor liner that includes a first annular panel section having a forward end, an aft end and a first cooling nugget at the forward end thereof, and a second annular panel section having a forward end, an aft end and a second cooling nugget at the forward end thereof. The second panel section is joined at its forward end to the aft end of the first panel section. One or more rows of cooling holes are located in the first cooling nugget, and one or more rows of cooling holes are located in the second cooling nugget. A group of dilution holes is located in the first panel section. The dilution holes are located at the aft end of the first panel section, immediately upstream of the second cooling nugget. Furthermore, each one of the dilution holes defines an aftmost edge, and all of the aftmost edges are axially aligned, even if the dilution holes have different hole diameters.
The present invention and its advantages over the prior art will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings.